1. Field of the Invention
This invention relates to spacecraft propulsion orbit control. Particularly, this invention relates to geosynchronous satellites employing electric propulsion systems for station keeping orbital control.
2. Description of the Related Art
Spacecraft such as communications satellites now commonly utilize electric propulsion for stationkeeping and other functions. Electric propulsion, such as ion thrusters utilize electrical power generated by the solar cells of the satellite to supply energy to a propellant to generate the propulsion. In general, ion thrusters possess a high specific impulse, making them extremely efficient requiring very little propellant for the thrust produced. Consequently, ion thrusters require relatively small amounts of a consumable propellant. This yields a significant advantage over conventional thrusters employing chemical constituents which react together to generate thrust.
Electric propulsion is the acceleration of propellants by electrical heating or electromagnetic forces. For example, ionized gas may be accelerated through an electric field across charged grids. The electrically accelerated particles can achieve speeds of approximately 30 km/second. The gas used is typically a noble gas, such as Xenon, for electromagnetic systems, and monopropellant for electrothermal systems. The principal advantage afforded by electric propulsion systems over conventional chemical propulsion systems is their very high efficiency. For example, with the same amount of fuel mass an ion propulsion system can achieve a final velocity as much as ten times higher than that obtainable with a chemical propulsion system.
Stationkeeping and momentum dumping on spacecraft, such as communications satellites, can require many hundreds of kilograms of propellant for conventional chemical thrusters. The use of electric propulsion systems has been applied to significantly reduce the mass of propellant needed. However, electric propulsion systems require additional electrical power to operate. Although additional electrical power can be generated from solar panels on-board the spacecraft, electric propulsion generally necessitates some additional solar array power, and thus additional solar array area, mass and cost.
U.S. Pat. No. 6,581,880 by Randolph et al., issued Jun. 24, 2003, discloses an electric propulsion device used to enable a stationkeeping satellite to track a prescribed stationkeeping Earth orbit. Electric propulsion propellant and electric power are throttled to vary the thrust and specific impulse of the electric propulsion device. A solar array provides electrical power during each Earth day cycle with excess power above that needed by the spacecraft stored by a battery. Software control manages the voltage, current and burn time to minimize propellant usage and impact to the system.
U.S. Pat. No. 6,341,749 by Ocampo, issued Jan. 29, 2002, discloses a method and apparatus for calculating an estimate of thrust vectors and burn times for an optimal two-burn orbit transfer from an inclined, eccentric initial orbit to a geostationary final orbit. A non-linear root finding algorithm is used to calculate the thrust vectors and burn times for the optimal two-burn orbit transfer. Thrust vectors and burn times are then computed for an optimal multi-segment orbit transfer from the initial orbit to the final orbit.
The system described in U.S. Pat. No. 6,581,880 takes advantage of times when there is excess solar array power to gain increased performance of the electric propulsion system. The system employs an electric propulsion system that can be throttled. To take full advantage of the system, it needs to have many different throttle settings, which adds hardware complexity, operational complexity, cost, and risk. In addition, the electrical power system for this and all conventional satellites employing electric propulsion are sized by a worst case day in the mission life using the electric propulsion system. Thus, in the prior art, the power system size is typically driven by the requirements of the electric propulsion system. As power system size continues to increase, strain is placed on the electrical power system in order to handle the additional power and current. Furthermore, it becomes more difficult to package the attitude control and structure system to support and control the larger solar array mass and area. In addition, the former system requires an electric propulsion system that can be throttled, with multiple throttle settings needed to take full advantage of the system. This feature adds hardware complexity, operational complexity, cost, and risk to the system.
In view of the foregoing, there is a need in the art for systems and methods to reduce the electrical power requirements when electric propulsion is employed. Furthermore, there is a need for in the art for attitude control systems and methods for that can be more easily packaged to support spacecraft solar arrays for spacecraft using electric propulsion without adding significant hardware complexity, operational complexity, cost, or risk. These and other needs are met by the present invention as detailed hereafter.